Low noise compressor and turbine for geared turbofan engine

ABSTRACT

A gas turbine engine has a fan section including a fan, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, and a gear reduction. The low pressure compressor and the low pressure turbine have a number of blades in each of at least one of a plurality of blade rows. The blades are rotatable at least some of the time at a rotational speed in operation. The number of blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: 5500≤(number of blades×rotational speed)/60≤10000, the rotational speed being in revolutions per minute.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.16/018,754, filed Jun. 26, 2018, which is a continuation of U.S. patentapplication Ser. No. 15/662,528, filed Jul. 28, 2017, which is acontinuation of U.S. patent application Ser. No. 15/270,027, filed Sep.20, 2016, which is a continuation of U.S. patent application Ser. No.15/014,363, filed Feb. 3, 2016, which is a continuation of U.S. patentapplication Ser. No. 14/967,478, filed Dec. 14, 2015, which is acontinuation-in-part of U.S. patent application Ser. No. 14/591,975,filed Jan. 8, 2015, which is a continuation-in-part of U.S. patentapplication Ser. No. 14/144,710, filed Dec. 31, 2013, which is acontinuation of U.S. patent application Ser. No. 14/016,436, filed Sep.3, 2013, now U.S. Pat. No. 8,714,913, issued May 6, 2014, which is acontinuation of U.S. patent application Ser. No. 13/630,276, filed Sep.28, 2012, now U.S. Pat. No. 8,632,301, issued Jan. 21, 2014.

BACKGROUND

This application relates to the design of a gas turbine engine rotorwhich can be operated to produce noise that is less sensitive to humanhearing.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor anddelivered downstream into a combustor section where it was mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving the turbine rotors to rotate.

Typically, there is a high pressure turbine rotor, and a low pressureturbine rotor. Each of the turbine rotors include a number of rows ofturbine blades which rotate with the rotor. Interspersed between therows of turbine blades are vanes.

The high pressure turbine rotor has typically driven a high pressurecompressor rotor, and the low pressure turbine rotor has typicallydriven a low pressure compressor rotor. Each of the compressor rotorsalso include a number of compressor blades which rotate with the rotors.There are also vanes interspersed between the rows of compressor blades.

The low pressure turbine or compressor can be a significant noisesource, as noise is produced by fluid dynamic interaction between theblade rows and the vane rows. These interactions produce tones at ablade passage frequency of each of the low pressure turbine rotors, thelow pressure compressor rotors, and their harmonics.

The noise can often be in a frequency range that is very sensitive tohumans. To mitigate this problem, in the past, a vane-to-blade ratio hasbeen controlled to be above a certain number. As an example, avane-to-blade ratio may be selected to be 1.5 or greater, to prevent afundamental blade passage tone from propagating to the far field. Thisis known as “cut-off.”

However, acoustically cut-off designs may come at the expense ofincreased weight and reduced aerodynamic efficiency. Stated another way,by limiting the designer to a particular vane to blade ratio, thedesigner may be restricted from selecting such a ratio based upon othercharacteristics of the intended engine.

Historically, the low pressure turbine has driven both a low pressurecompressor section and a fan section. More recently, a gear reductionhas been provided such that the fan and low pressure compressor can bedriven at distinct speeds.

SUMMARY

In a featured embodiment, a gas turbine engine has a fan, a compressorsection having a low pressure portion and a high pressure portion, acombustor section, and a turbine having a first turbine rotor. The firstturbine rotor drives the fan. A gear reduction effects a reduction inthe speed of the fan relative to a speed of the first turbine rotor.Each of the compressor rotor and the first turbine rotor includes anumber of blades in each of a plurality of rows. The blades operate atleast some of the time at a rotational speed. The number of blades andthe rotational speed are such that the following formula holds true forat least one of the blade rows of the first turbine rotor and/or thecompressor rotor: (number of blades×rotational speed)/60≥5500. Therotational speed is an approach speed in revolutions per minute.

These and other features of this application will be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2 shows another embodiment.

FIG. 3 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown), or an intermediate spool,among other systems or features. The fan section 22 drives air along abypass flowpath B in a bypass duct defined within a nacelle 15, whilethe compressor section 24 drives air along a core flowpath C forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The terms “low” and “high” as applied to speed or pressure for thespools, compressors and turbines are of course relative to each other.That is, the low speed spool operates at a lower speed than the highspeed spool, and the low pressure sections operate at lower pressurethan the high pressures sections.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. In some embodiments, the bypass ratio isless than about thirty (30), or more narrowly less than about twenty(20). In embodiments, the gear reduction ratio is less than about 5.0,or less than about 4.0. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a geared architectureengine and that the present invention is applicable to other gas turbineengines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient°R)/(518.7)°R]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second.

The use of the gear reduction between the low speed spool and the fanallows an increase of speed to the low pressure compressor. In the past,the speed of the low pressure turbine and compressor has been somewhatlimited in that the fan speed cannot be unduly large. The maximum fanspeed is at its outer tip, and in larger engines, the fan diameter ismuch larger than it may be in smaller power engines. However, the use ofthe gear reduction has freed the designer from limitation on the lowpressure turbine and compressor speeds caused by a desire to not haveunduly high fan speeds.

It has been discovered that a careful design between the number ofrotating blades, and the rotational speed of the low pressure turbinecan be selected to result in noise frequencies that are less sensitiveto human hearing. The same is true for the low pressure compressor 44.

A formula has been developed as follows:

(blade count×rotational speed)/60 sec≥5500 Hz.

That is, the number of rotating blades in any low pressure turbinestage, multiplied by the rotational speed of the low pressure turbine 46(in revolutions per minute), divided by 60 sec should be greater than orequal to about 5500 Hz. The same holds true for the low pressurecompressor stages. More narrowly, the amounts should be greater than orequal to about 6000 Hz. In embodiments, the amount is less than or equalto about 10000 Hz, or more narrowly less than or equal to about 7000 Hz.A worker of ordinary skill in the art would recognize that the 60 secfactor is to change revolutions per minute to Hertz, or revolutions perone second. For the purposes of this disclosure, the term “about” means±3% of the respective quantity unless otherwise disclosed.

The operational speed of the low pressure turbine 46 and low pressurecompressor 44 as utilized in the formula should correspond to the engineoperating conditions at each noise certification point defined in Part36 or the Federal Airworthiness Regulations. More particularly, therotational speed may be taken as an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations. Forpurposes of this application and its claims, the term “approach speed”equates to this certification point. In other embodiments, the aboveformula results in a number that is less than or equal to about 7000 Hzat approach speed.

It is envisioned that all of the rows in the low pressure turbine 46meet the above formula. However, this application may also extend to lowpressure turbines wherein the majority of the blade rows, or at leasthalf of the blade rows, in the low pressure turbine meet the aboveformula, but perhaps some may not. By implication at least one, or lessthan half, of the rows meet the formula. The same is true for lowpressure compressors, wherein all of the rows in the low pressurecompressor 44 would meet the above formula. However, the application mayextend to low pressure compressors wherein only the majority of theblade rows, or at least half of the blade rows, in the low pressurecompressor meet the above formula, but some perhaps may not. Of course,by implication the formula may be true for at least some of the turbinerows but no compressor rows. In some cases, only one row of the lowpressure turbine and/or low pressure compressor may meet the formula.Also, the formula may apply to at least some compressor rows, but no rowin the turbine meets the formula.

This will result in operational noise that would be less sensitive tohuman hearing.

In embodiments, it may be that the formula can result in a range ofgreater than or equal to 5500 Hz, and moving higher. Thus, by carefullydesigning the number of blades and controlling the operational speed ofthe low pressure turbine 46 (and a worker of ordinary skill in the artwould recognize how to control this speed) one can assure that the noisefrequencies produced by the low pressure turbine are of less concern tohumans.

The same holds true for designing the number of blades and controllingthe speed of the low pressure compressor 44. Again, a worker of ordinaryskill in the art would recognize how to control the speed.

In embodiments, it may be only the low pressure turbine rotor 46, or thelow pressure compressor rotor 44 which is designed to meet the meet theabove formula. On the other hand, it is also possible to ensure thatboth the low pressure turbine 46 and low pressure compressor 44 meet theabove formula.

This invention is most applicable to jet engines rated to produce 15,000pounds of thrust or more. In this thrust range, prior art jet engineshave typically had frequency ranges of about 4000 hertz. Thus, the noiseproblems as mentioned above have existed.

Lower thrust engines (<15,000 pounds) may have operated under conditionsthat sometimes passed above the 4000 Hz number, and even approached 6000Hz, however, this has not been in combination with the gearedarchitecture, nor in the higher powered engines which have the largerfans, and thus the greater limitations on low pressure turbine or lowpressure compressor speed.

FIG. 2 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 3 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

The FIGS. 2 and 3 engines may be utilized with the speed and bladefeatures disclosed above.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

what is claimed is:
 1. A gas turbine engine comprising: a fan sectionincluding a fan, and a low fan pressure ratio of less than 1.45, whereinthe low fan pressure ratio is measured across a fan blade alone; acompressor section including a low pressure compressor and a highpressure compressor; wherein the fan delivers air into a bypass ductdefined within a nacelle, and a portion of air into the compressorsection, with a bypass ratio defined as the volume of air delivered intothe bypass duct compared to the volume of air delivered into thecompressor section, and the bypass ratio being greater than 10; aturbine section including a low pressure turbine and a high pressureturbine; wherein the low pressure turbine includes an inlet, an outletand a pressure ratio of greater than 5, the pressure ratio beingpressure measured prior to the inlet as related to pressure at theoutlet prior to an exhaust nozzle; a gear reduction including anepicyclic gear train, wherein the gear reduction effects a reduction inthe speed of the fan relative to a speed of the low pressure turbine,the epicyclic gear train having a gear reduction ratio of greater than2.5:1; wherein each of the low pressure compressor and the low pressureturbine includes a number of blades in each of a plurality of bladerows, the number of blades rotatable at least some of the time at arotational speed in operation, and the number of blades and therotational speed being such that the following formula holds true for atleast one of the blade rows of the low pressure compressor:5500 Hz≤(number of blades×rotational speed)/60≤10000 Hz, the rotationalspeed being an approach speed in revolutions per minute, taken at anapproach certification point as defined in Part 36 of the FederalAirworthiness Regulations; and the following formula holds true for atleast one of the blade rows of the low pressure turbine:(number of blades×rotational speed)/60<10000 Hz, the rotational speedbeing an approach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations; and wherein the engine is rated to produce 15,000 pounds ofthrust or more.
 2. The gas turbine engine as set forth in claim 1,wherein the formula results in a number less than or equal to 10000 forat least a plurality of the blade rows of the low pressure compressor.3. The gas turbine engine as set forth in claim 2, wherein the formularesults in a number greater than or equal to 5500 for at least half ofthe blade rows of the low pressure compressor.
 4. The gas turbine engineas set forth in claim 3, wherein the formula results in a number greaterthan or equal to 5500 for at least one of the blade rows of the lowpressure turbine.
 5. The gas turbine engine as set forth in claim 4,wherein the formula results in a number greater than or equal to 6000for at least one of the blade rows of the low pressure compressor. 6.The gas turbine engine as set forth in claim 5, wherein the formularesults in a number less than or equal to 10000 for at least a pluralityof the blade rows of the low pressure turbine.
 7. The gas turbine engineas set forth in claim 6, wherein the formula results in a number greaterthan or equal to 6000 for at least half of the blade rows of the lowpressure compressor.
 8. The gas turbine engine as set forth in claim 7,wherein the formula results in a number greater than or equal to 6000for at least half of the blade rows of the low pressure turbine.
 9. Thegas turbine engine as set forth in claim 8, wherein the formula resultsin a number greater than or equal to 6000 for three of the blade rows ofthe low pressure turbine.
 10. The gas turbine engine as set forth inclaim 9, wherein the formula results in a number less than or equal to10000 for at least a majority of the blade rows of the low pressureturbine.
 11. The gas turbine engine as set forth in claim 10, whereinthe formula results in a number less than or equal to 10000 for at leasta majority of the blade rows of the low pressure compressor.
 12. The gasturbine engine as set forth in claim 11, further comprising: a coreflowpath and a mid-turbine frame arranged between the low pressureturbine and the high pressure turbine, the mid-turbine frame havingairfoils positioned in the core flowpath, and the mid-turbine framesupporting at least one bearing system; wherein the low pressurecompressor includes three stages, the low pressure turbine includes agreater number of stages than the low pressure compressor, the highpressure turbine includes two stages, and the high pressure compressorincludes a greater number of stages than the low pressure turbine; andwherein the fan has a low corrected fan tip speed of less than 1150ft/second.
 13. The gas turbine engine as set forth in claim 12, whereinthe gear reduction is a planetary gear system.
 14. A gas turbine enginecomprising: a fan section including a fan, and a low fan pressure ratioof less than 1.45, wherein the low fan pressure ratio is measured acrossa fan blade alone; a compressor section including a low pressurecompressor and a high pressure compressor; wherein the fan delivers airinto a bypass duct defined within a nacelle, and a portion of air intothe compressor section, with a bypass ratio defined as the volume of airdelivered into the bypass duct compared to the volume of air deliveredinto the compressor section, and the bypass ratio being greater than 10;a turbine section including a low pressure turbine and a high pressureturbine; wherein the low pressure turbine includes an inlet, an outletand a pressure ratio of greater than 5, the pressure ratio beingpressure measured prior to the inlet as related to pressure at theoutlet prior to an exhaust nozzle; a gear reduction including anepicyclic gear train, wherein the gear reduction is positionedintermediate the low pressure compressor and a shaft driven by the lowpressure turbine such that a fan rotor of the fan section and the lowpressure compressor are rotatable at a common speed in operation, theepicyclic gear train having a gear reduction ratio of greater than2.5:1; wherein each of the low pressure compressor and the low pressureturbine includes a number of blades in each of a plurality of bladerows, the number of blades rotatable at least some of the time at arotational speed in operation, and the number of blades and therotational speed being such that the following formula holds true for atleast one of the blade rows of the low pressure compressor:5500 Hz≤(number of blades x rotational speed)/60≤10000 Hz, therotational speed being an approach speed in revolutions per minute,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations; and the following formula holds truefor at least one of the blade rows of the low pressure turbine:(number of blades×rotational speed)/60≤10000 Hz, the rotational speedbeing an approach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations; and wherein the engine is rated to produce 15,000 pounds ofthrust or more.
 15. The gas turbine engine as set forth in claim 14,wherein the formula results in a number less than or equal to 10000 forat least a plurality of the blade rows of the low pressure compressor.16. The gas turbine engine as set forth in claim 15, wherein the formularesults in a number greater than or equal to 5500 for at least one ofthe blade rows of the low pressure turbine.
 17. The gas turbine engineas set forth in claim 16, wherein the formula results in a numbergreater than or equal to 6000 for at least one of the blade rows of thelow pressure compressor and for at least one of the blade rows of thelow pressure turbine.
 18. The gas turbine engine as set forth in claim17, wherein the formula results in a number greater than or equal to6000 for at least half of the blade rows of the low pressure compressor.19. The gas turbine engine as set forth in claim 18, wherein the formularesults in a number greater than or equal to 6000 for at least half ofthe blade rows of the low pressure turbine.
 20. The gas turbine engineas set forth in claim 19, wherein the formula results in a number lessthan or equal to 10000 for at least a plurality of the blade rows of thelow pressure turbine.
 21. The gas turbine engine as set forth in claim20, wherein the formula results in a number less than or equal to 7000for only a majority of the blade rows of the low pressure compressor andfor only a majority of the blade rows of the low pressure turbine. 22.The gas turbine engine as set forth in claim 20, wherein the formularesults in a number less than or equal to 10000 for at least a majorityof the blade rows of the low pressure compressor and for at least amajority of the blade rows of the low pressure turbine.
 23. The gasturbine engine as set forth in claim 22, wherein the formula results ina number less than or equal to 7000 for at least a majority of the bladerows of the low pressure compressor and for at least a majority of theblade rows of the low pressure turbine, and the formula results in anumber greater than or equal to 6000 for at least a majority of theblade rows of the low pressure compressor and for at least a majority ofthe blade rows of the low pressure turbine.
 24. The gas turbine engineas set forth in claim 23, further comprising: a core flowpath and amid-turbine frame arranged between the low pressure turbine and the highpressure turbine, the mid-turbine frame having airfoils positioned inthe core flowpath, and the mid-turbine frame supporting at least onebearing system; wherein the low pressure compressor includes threestages, the low pressure turbine includes five stages, the high pressureturbine includes two stages, and the high pressure compressor includeseight stages; and wherein the fan has a low corrected fan tip speed ofless than 1150 ft/second.
 25. A gas turbine engine comprising: a fansection including a fan, and a low fan pressure ratio of less than 1.45,wherein the low fan pressure ratio is measured across a fan blade alone;a compressor section including a low pressure compressor, anintermediate pressure compressor and a high pressure compressor; whereinthe fan delivers air into a bypass duct defined within a nacelle, and aportion of air into the compressor section, with a bypass ratio definedas the volume of air delivered into the bypass duct compared to thevolume of air delivered into the compressor section, and the bypassratio being greater than 10; a turbine section including a low pressureturbine, an intermediate pressure turbine and a high pressure turbine,the intermediate pressure turbine driving the intermediate pressurecompressor, and the high pressure turbine driving the high pressurecompressor; wherein the low pressure turbine includes an inlet, anoutlet and a pressure ratio of greater than 5, the pressure ratio beingpressure measured prior to the inlet as related to pressure at theoutlet prior to an exhaust nozzle; a gear reduction including anepicyclic gear train, wherein the gear reduction effects a reduction inthe speed of the fan relative to a speed of the low pressure turbine,the epicyclic gear train having a gear reduction ratio of greater than2.5:1; wherein each of the low pressure compressor and the low pressureturbine includes a number of blades in each of a plurality of bladerows, the number of blades rotatable at least some of the time at arotational speed in operation, and the number of blades and therotational speed being such that the following formula holds true for atleast one of the blade rows of the low pressure compressor:6000 Hz≤(number of blades x rotational speed)/60≤10000 Hz, therotational speed being an approach speed in revolutions per minute,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations; and the following formula holds truefor at least one of the blade rows of the low pressure turbine:(number of blades×rotational speed)/60≤10000 Hz, the rotational speedbeing an approach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations; and wherein the engine is rated to produce 15,000 pounds ofthrust or more.
 26. The gas turbine engine as set forth in claim 25,wherein the formula results in a number less than or equal to 10000 forat least a plurality of the blade rows of the low pressure compressorand for at least a plurality of the blade rows of the low pressureturbine.
 27. The gas turbine engine as set forth in claim 26, whereinthe formula results in a number greater than or equal to 5500 for atleast half of the blade rows of the low pressure compressor and for atleast half of the blade rows of the low pressure turbine.
 28. The gasturbine engine as set forth in claim 27, wherein the formula results ina number less than or equal to 7000 for at least a majority of the bladerows of the low pressure compressor and for at least a majority of theblade rows of the low pressure turbine.
 29. The gas turbine engine asset forth in claim 28, wherein the formula results in a number greaterthan or equal to 6000 for at least a majority of the blade rows of thelow pressure compressor and for at least a majority of the blade rows ofthe low pressure turbine.
 30. The gas turbine engine as set forth inclaim 29, further comprising: a core flowpath and a mid-turbine framearranged between the low pressure turbine and the high pressure turbine,the mid-turbine frame having airfoils positioned in the core flowpath,and the mid-turbine frame supporting at least one bearing system;wherein the gear reduction is a planetary gear system, and the gearreduction is positioned intermediate the low pressure compressor and ashaft driven by the low pressure turbine such that a fan rotor of thefan section and the low pressure compressor are rotatable at a commonspeed in operation; wherein the low pressure compressor includes threestages, the low pressure turbine includes five stages, the high pressureturbine includes two stages, and the high pressure compressor includeseight stages; and wherein the fan has a low corrected fan tip speed ofless than 1150 ft/second.